(1) Field of the Invention
The present invention relates to an adaptively-twistable blade and to an aircraft provided with such a blade, and more particularly but not exclusively to a blade for a rotorcraft lift rotor.
(2) Description of Related Art
Conventionally, a blade extends longitudinally from a first end for fastening to a rotary hub of a rotor towards a second end that is referred to as a “free” end.
The blade also extends transversely from a leading edge towards a trailing edge. The blade includes in particular an outer covering having a first skin for its suction side and referred to for convenience as its “suction side” skin, and a second skin for its pressure side and referred to for convenience as its “pressure side” skin.
A blade of a main lift rotor of a rotorcraft exerts lift during rotary motion of said main rotor, serving to support the rotorcraft in the air and possibly also to propel it. Depending on the pitch angle of the blade as controlled by a pilot or an autopilot system, the lift developed by the blade can be made greater or smaller. The aerodynamic angle of incidence of each aerodynamic profile of the blade depends on the pitch angle of the blade. An aerodynamic profile is referred to simply as “profile” for convenience.
However, starting from a threshold angle of incidence, the air stream over a given profile is observed to separate therefrom and thus from a given section of the blade. By way of example, this separation may occur at the leading edge or at the trailing edge of the profile, or indeed over the suction side of the blade at a distance lying in the range 50% to 60% along the chord of said profile from the leading edge. Such separation causes the blade to stall, i.e. gives rise to a sudden drop in its lift if the phenomenon propagates and persists over a zone lying between two profiles defining a critical surface along the span of the blade. Furthermore, separation of the air stream gives rise to turbulence, which can lead to vibration and to an increase in the drag coefficient of the blade.
Likewise, because of the different local speeds along the span of the blade, the values of lift and drag forces at any point increase ongoing from the root towards the end of a blade that is straight. This leads to non-uniform distribution of lift and drag, and to considerable moments at the root. These effects are harmful to the mechanical integrity of the blade and to the flight qualities of the rotary wing.
To limit those effects, one solution consists in twisting the blade geometrically. Twisting has the effect of limiting separation of the boundary layer over the entire span of the blade, and thus of shifting the resultant of the lift and drag forces towards the root of the blade so as to reduce the resulting movements at the root. It should be observed that the geometrical twist of a blade may be defined by the angle formed between the chord of each profile of a section of the blade relative to the reference plane of the blade. Sometimes each profile of the blade is twisted relative to the pitch variation axis of the blade through an angle measured relative to such a reference plane.
For any given blade path, it can be understood that twisting has a direct influence on the aerodynamic angle of incidence of each profile. Under such conditions, the term “twisting relationship” designates how said twist angles vary along the span of the blade.
For a given blade, the twisting relationship of that blade does not vary. The twisting relationship is the result of an acceptable compromise for optimizing the operation of the rotor over the entire flight envelope of the aircraft.
Specifically, it is found for example that a small amplitude of twist over the entire span of the blade serves to minimize the power consumed by the lift rotor of a rotorcraft in forward flight. Conversely, a large amplitude of twist over the entire span of the blade serves to minimize the power consumed by the lift rotor of a rotorcraft in hovering flight, but is penalizing during forward flight. It should be understood that a “small” amplitude is used to designate an amplitude of twist variation lying in the range 4° to 8°, for example, whereas a “large” amplitude is used to designate an amplitude lying in the range 16° to 20°, for example.
Thus, a twist amplitude lying between those small and large amplitudes represents a compromise in terms of power consumption between stages of forward flight and stages of hovering flight.
In order to avoid making such a compromise, proposals have been made to modify the twist of a blade actively, at least locally. Thus, by adapting the twist of the blade to specific flight configurations, the performance of a rotorcraft can be improved considerably in terms of transportable payload and/or cruising speed. Its environmental impact can also be reduced by lowering fuel consumption, for performance that is equivalent or even improved compared with a rotorcraft having blades of non-adaptable twist angle.
In order to twist a blade, one technique consists in deforming a structure of the blade in torsion. Such deformation may be performed with the help of an actuator delivering a drive force for twisting the blade. Such actuators may be associated with installing a blade structure made of composite materials using specific draping of fibers.
In one embodiment, a mass is moved along the chord direction at the free end of a blade in order to shift the center of gravity of the blade. Such a shift gives rise to a torsional moment on the blade about the pitch axis of the blade, for example, resulting from the combination of conical shape and centrifugal force. The mass may be moved with the help of an actuator. In addition, the torsional moment exerted on a blade section may be transmitted to another blade section via a torsion bar.
Documents US 2013/0062456, US 2012/0153073, and U.S. Pat. No. 5,505,589 describe solutions of that type.
The use of an actuator is advantageous. Nevertheless, it can be difficult to install an actuator within a blade, and in particular within a rotorcraft blade.
The space available for installing it is small, thereby putting a limit on the dimensions of the actuator that is to be used.
In addition, the actuator is subjected to high levels of mechanical stress (centrifugal force, vibration).
If an electric actuator is installed, the electric actuator also needs to be powered electrically and that can be difficult to achieve in a rotating frame of reference. Transferring electrical power from a stationary frame of reference corresponding to the airframe of the rotorcraft to a rotating frame of reference corresponding to the rotor can be constraining. The level of difficulty encountered depends on the voltage or the current of the electricity to be transferred.
In order to power an actuator of electromechanical type having a power of a few kilowatts, a set of slip rings can be used to transfer electricity from a generator situated in the rotorcraft to an actuator situated in a blade. Requirements in terms of performance and reliability make such a set of slip rings expensive.
An actuator of piezoelectric type requires current that is low compared with an electromechanical actuator, but it requires electricity to be delivered at a much higher voltage. This constraint complicates the set of slip rings and significantly increases its costs.
The use of actuators requiring high levels of electrical power can thus give rise to problems of size and to installation costs that are high.
Unfortunately, the structure of a blade may present a relatively high degree of stiffness in torsion in order to present sound dynamic behavior. Modifying such a structure in twisting tends to require powerful actuators to be put into place.
The use of such actuators for high powers thus involves technical solutions that are complex and expensive in order to ensure that such systems operate properly with a high level of reliability.
In another embodiment, the structure of a blade made of composite material includes fibers presenting a particular orientation for the purpose of causing the blade to twist under the effect of centrifugal force. Traction exerted on those fibers by centrifugal force generates torsion of the blade. Such fibers may be arranged in the suction side and pressure side skins of the blade, or in a strip that is present within the blade, for example.
Documents US 2012/0153073, FR 2 737 465, and FR 2 956 856 describe systems making use of this principle.
Document U.S. Pat. No. 7,037,076 suggests using a shape memory alloy for causing the end of a blade to turn.
Document FR 2 374 212 discloses a blade having two spars fastened to a pitch control shaft so that a change in the pitch of the blade causes the spars to twist.
Other techniques are remote from the invention and seek to modify the profiles of the blades instead of moving them.
Thus, Documents WO 2007/079855, WO 2004/069651, US 2006/0038058, and EP 1 144 248 propose modifying the profiles of sections of a blade with the help of piezoelectric actuators. Such deformation is sometimes referred to as “morphing”.
Documents FR 2 924 681, U.S. Pat. No. 5,004,189, US 2008/0145220, and U.S. Pat. No. 7,677,868 make use of electromechanical actuators.
Document U.S. Pat. No. 7,837,144 provides for a pneumatic actuator.
Documents WO 2010/043645, WO 2010/023286, WO 2010/023278, and WO 2009/056136 relate to the remote field of wind turbine blades.